1. Field of the Invention
The present invention relates to a combustor for a gas turbine engine and, more particularly, to a combustor for a gas turbine engine having a compact design which operates efficiently at inlet air flows having a high subsonic Mach Number.
2. Description of Related Art
Advanced aircraft gas turbine engine technology requirements are driving the combustors therein to be shorter in length, have higher performance levels over wider operating ranges, and produce lower exhaust pollutant emission levels. To achieve this, such combustion system designs have become considerably more complex and costly compared with current technology combustion systems. For example, combustors having multiple annular dome configurations, such as the double annular combustor disclosed in U.S. Pat. No. 5,197,278 to Sabla et al. and the triple annular combustor disclosed in U.S. Pat. No. 5,323,604 to Eckstedt et al., have evolved which evidence the complexity of such advanced combustion systems. While these combustor designs are able to perform in accordance with their stated objectives, they are complex in design, generally requiring complicated fuel staging modes to maintain stability over their intended operating range and involve a substantial number of sub-assemblies for pre-mixing air and fuel in order to achieve a desired reduction in gas turbine emissions. Further, such combustor designs continue to be concerned with auto-ignition within the mixers and generally have involved an overall increase in the height and weight of the combustion system.
Accordingly, it would be desirable for a gas turbine engine combustor to be developed which avoids the problems associated with multiple annular combustors. Further, it would be desirable if the combustor could be more compact in both height and length, as well as substantially reduce combustion system pressure loss.